Thermal barrier coating resistant to penetration by environmental contaminants

ABSTRACT

A turbine engine component includes an electron beam-physical vapor deposition thermal barrier coating covering at least a portion of a substrate. The thermal barrier coating includes an inner layer having a columnar-grained microstructure with inter-columnar gap porosity. The inner layer includes a stabilized ceramic material. The thermal barrier coating also includes a substantially non-porous outer layer, covering the inner layer and including the stabilized ceramic material. The outer layer is deposited with continuous line-of-sight exposure to the vapor source under oxygen deficient conditions. The outer layer may further comprise a dopant oxide that is more readily reducible than the stabilized ceramic material. During deposition, the outer layer may also have an oxygen deficient stoichiometry with respect to the inner layer. Oxygen stoichiometry in the outer layer may be restored by exposure of the coated component to an oxidizing environment.

TECHNICAL FIELD

The present invention relates to thermal barrier-coated turbine enginecomponents that function in high temperature environments. Moreparticularly, the present invention relates to coatings for turbineengine components to improve resistance to heat and contaminants thatare commonly part of high-temperature combustion gas environments.

BACKGROUND

Turbine engines are used as the primary power source for various kindsof aircrafts. The engines are also auxiliary power sources that driveair compressors, hydraulic pumps, and industrial gas turbine (IGT) powergeneration. Further, the power from turbine engines is used forstationary power supplies such as backup electrical generators.

Most turbine engines generally follow the same basic power generationprocedure. Compressed air is mixed with fuel and burned, and theexpanding hot combustion gases are directed against stationary turbinevanes in the engine. The vanes turn the high velocity gas flow partiallysideways to impinge on the turbine blades mounted on a rotatable turbinedisk. The force of the impinging gas causes the turbine disk to spin athigh speed. Jet propulsion engines use the power created by the rotatingturbine disk to draw more air into the engine and the high velocitycombustion gas is passed out of the gas turbine aft end to createforward thrust. Other engines use this power to operate one or morepropellers, electrical generators, or other devices.

Many turbine engine and aeroengine components such as blades, guidevanes, combustor cans, and so forth are formed from a superalloy, andare often coated with a thermal barrier coating to extend the componentlife. Since a temperature gradient is produced across the thermalbarrier coating during engine operation, the engine component functionsat a reduced temperature with respect to the operating environment. Inaddition to providing a thermal barrier, if the coating material has athermal expansion coefficient that differs from that of the underlyingcomponent material, the coating material typically is processed to haveporosity that provides high in-plane compliance to accommodate a thermalexpansion mismatch.

Both the protective properties and the in-plane compliance for thethermal barrier coating may be adversely affected if the enginecomponent is exposed to some types of environmental contaminants. Oneclass of contaminants that may potentially reduce a thermal barriercoating's protective and compliance characteristics includes dust,comprising oxides of calcium, magnesium, aluminum, silicon, and mixturesthereof, which are commonly referred to as CMAS. Another class ofcontaminants that can wick into porous thermal barrier coatings ismolten sulfate salts, such as sodium sulfate, which is a constituent ofsea salt. Molten CMAS and sulfate salts may penetrate the pores orchannels in a thermal barrier coating. Upon cooling, the penetratedcontaminates solidify and thereby reduce the coating's in-planecompliance. Cracking, fragmentation, and spalling in the thermal barriercoating may result from the reduced ability to tolerate compressivestrain.

Hence, there is a need for a substrate coating that has thermal barrierproperties, high in-plane compliance and is resistant to contaminationand penetration from environmental contaminants such as CMAS that existin a high temperature system. There is a further need for efficientmethods for manufacturing a component that includes such a coating.

BRIEF SUMMARY

The present invention provides a turbine engine component, comprising asubstrate, and a thermal barrier coating covering at least a portion ofthe substrate. The thermal barrier coating includes an inner layerhaving a columnar-grained microstructure with inter-columnar gapporosity. The inner layer includes a stabilized ceramic material. Thethermal barrier coating also includes a substantially non-porousstabilized ceramic outer layer, covering the inner layer. The stabilizedceramic outer layer may comprise a dopant oxide that is more readilyreducible than the stabilized ceramic material. The stabilized ceramicouter layer may also be deposited with an oxygen deficient stoichiometrywith respect to the inner layer.

The present invention also provides a first method of protecting aturbine engine component from heat and environmental contaminants. Themethod includes the steps of depositing a thermal barrier coating innerlayer, including a stabilized ceramic material having a columnarmicrostructure with inter-columnar gap porosity, onto at least a regionof a component surface by rotating or otherwise intermittentlypositioning the component surface region requiring coating within theline-of-sight of an electron beam-physical vapor deposition vaporsource, and subsequently depositing a substantially non-porous thermalbarrier coating outer layer. The outer layer of the thermal barriercoating comprises a stabilized ceramic material which is deposited withan oxygen deficient stoichiometry. The outer layer of thermal barriercoating may further comprise a dopant oxide that is more readilyreducible than the stabilized ceramic material. Surfaces requiring thedense layer of thermal barrier coating may be stationary or continuouslyexposed to deposition from the line-of-sight electron beam-physicalvapor deposition vapor source. The outer layer of the thermal barriercoating is substantially free of interconnected porosity and has ahigher density than the inner layer.

The present invention also provides a second method of protecting aturbine engine component from heat and environmental contaminants. Themethod comprises the steps of depositing a thermal barrier coating innerlayer, comprising a stabilized ceramic material having a columnarmicrostructure with inter-columnar gap porosity, onto at least a regionof a component surface, by depositing a ceramic material onto thesubstrate in a deposition chamber while bleeding oxygen into thedeposition chamber, and forming a substantially non-porous thermalbarrier coating outer layer onto the inner layer by depositing theceramic material onto the inner layer without bleeding oxygen into thedeposition chamber.

Other independent features and advantages of the preferred turbineengine component and protecting method will become apparent from thefollowing detailed description, taken in conjunction with theaccompanying drawings which illustrate, by way of example, theprinciples of the invention.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a perspective view of a blade that is exemplary of the typesthat are used in turbine engines;

FIG. 2 is a cross-sectional view of a substrate coated with a densifiedthermal barrier coating, illustrating the architecture of an exemplarycoating system according to an embodiment of the present invention; and

FIG. 3 is a flow chart illustrating production steps for coating asubstrate with a densified thermal barrier coating according to anembodiment of the present invention.

DETAILED DESCRIPTION OF A PREFERRED EMBODIMENT

The following detailed description of the invention is merely exemplaryin nature and is not intended to limit the invention or the applicationand uses of the invention. Furthermore, there is no intention to bebound by any theory presented in the preceding background of theinvention or the following detailed description of the invention.

The present invention includes a coating for a variety of substrates,including gas turbine and aeroengine components. The coating has boththermal barrier properties and high in-plane compliance. Further, thecoating includes a thin densified outermost region that is resistant topenetration from environmental contaminants such as molten CMAS depositsand sulfate salts. The densified outermost region keeps contaminantsfrom wicking into the coating and thereby reducing the coating'sin-plane compliance. As a result, contaminants are initially kept on thecoating outer surface, and can be removed by evaporation into theflowing combustion gases. For example, at high temperatures silica andalumina may react with water vapor that is present in combustion gases,and evaporate from the deposits as hydroxides.

FIG. 1 illustrates a superalloy blade 150 that is exemplary of the typesthat are used in turbine engines, although turbine blades commonly havedifferent shapes, dimensions and sizes depending on gas turbine enginemodels and applications. Nickel-based superalloys are just one class ofmaterials that are commonly used to manufacture turbine engine blades.The illustrated blade 150 has an airfoil portion 152 including apressure surface 153, an attachment or root portion 154, a leading edge158 including a blade tip 155, and a platform 156. The blade 150 may beformed with a non-illustrated outer shroud attached to the tip 155. Theblade 150 may have non-illustrated internal air-cooling passages thatremove heat from the turbine airfoil. After the internal air hasabsorbed heat from the superalloy, the air is discharged into thecombustion gas flowpath through passages 159 in the airfoil wall.

As mentioned previously, the densified thermal barrier coating of thepresent invention can be tailored to fit a blade's specific needs, whichdepend in part on the blade component where degradation may occur. Forexample, the densified layer of the thermal barrier coating may beselectively applied to blade surfaces that are potentially exposed todeposition of environmental contaminants. In another exemplaryembodiment, the densified layer of the thermal barrier coating isthicker at particular locations that are most likely to beaerodynamically impacted by CMAS dust particles. For example, duringturbine engine operation particulate contaminants, such as CMASparticles, preferentially deposit on the leading edge 158 and theairfoil pressure surface 153 of turbine blade 150. The densified thermalbarrier coating layer may be applied solely or primarily to these bladeareas where deposits are anticipated. On blade areas where moltencontaminant deposits will not be formed, the porous inner layer of theTBC is viable by itself without an overlying dense layer.

It is also emphasized again that turbine blades are just one example ofthe type of turbine components to which the densified thermal barriercoating of the present invention may be applied. Vanes, shrouds, andother turbine components can be coated in the same manner.

Turning now to FIG. 2, a cross-sectionial view of a substrate 10 coatedwith an exemplary densified thermal barrier coating 40 is illustrated.The thermal barrier coating 40 overlies the substrate 10 and anyintermediate layers, and is formed of a ceramic material. In the contextof a turbine engine airfoil, the thermal barrier coating 40 insulatesthe substrate 10 from the high temperature exhaust gas passing over theairfoil surface during engine operation. The thermal barrier coating 40is a two-layer structure that includes an inner layer 42, and a muchthinner outer layer 44. The inner layer 42 may be any acceptablematerial, and an exemplary material has a columnar-grained structure.The columnar-grained structure is porous having side-by-side columnsthat grow substantially perpendicular to the substrate surface and anyintermediate layers. The columns are spaced apart to some degree, andthe columns themselves are somewhat porous. The spaces and pores reducethe effects of stresses and strains induced when the protected articleis repeatedly heated and cooled during service. Exemplary materials forthe inner layer 42 include stabilized tetragonal and cubic zirconia andstabilized tetragonal and cubic hafnia, and mixtures thereof. Apreferred example is about 7 weight % yttria-stabilized tetragonalzirconia.

The inner layer 42 can also be formed as a columnar-grained structurednanolaminate. For example, U.S Pat. No. 6,482,537, which is incorporatedherein by reference, discloses a thermal barrier coating that includes acolumnar grained ceramic layer applied to an aluminide or MCrAlY bondcoat by an electron beam-physical vapor deposition process. The ceramiclayer includes a plurality of layers of stabilized zirconia, with theinterfaces between the layers decorated with a secondary ceramicconstituent, which includes particles selected from the group consistingof tantala and alumina. The multilayered ceramic coating includes atleast 85% yttria stabilized zirconia, and between 1 and 15% secondaryconstituent.

Because the thermal barrier coating inner layer 42 is porous andincludes spaces between the column structures, the inner layer 42 issomewhat vulnerable to attack by molten environmental contaminants suchas CMAS, which may wick into the microstructure. To protect the innerlayer 42 from penetration by molten contaminants, the outer layer 44 ofthe thermal barrier coating is a densified oxide layer. The outer layermay be simply a modification of the inner layer 42 in terms ofstructure, and does not exhibit the porosity of the inner layer 42. Theouter and inner layers may have essentially the same chemical makeup. Inother words, the outer layer 44 may be formed using the same compoundsas the inner layer 42, only using a process that creates a densematerial when compared with the porous columnar inner layer 42. Theouter layer 44 may also be doped with additional oxide constituents thatare more readily reduced than the stabilized ceramic material, or withan oxide that enhances diffusion. The concentration of the oxide dopantsin the outer layer 44 ranges between about 0.5% and about 20%, and suchoxides are in addition to the yttria or other stabilizing oxide.Exemplary oxide constituents include tantala, niobia, and alumina, whichenhance surface diffusion and densification of the thermal barriercoating.

The thermal barrier coating outer layer 44 is very thin with respect tothe inner layer 42. In an exemplary embodiment, the thickness of thecolumnar inner layer 42 is between about 1 and about 10 mils (about 25to about 250 μm), while the dense outer layer 44 is less than about 1 μm(about 25 μm), and preferably no greater than 0.2 mils (about 5 μm) inthickness. Also, the outer layer 44 is substantially free ofinterconnected porosity that could wick molten CMAS or sulfate saltdeposits into the thicker compliant thermal barrier coating layer 42.

In the exemplary structure illustrated in FIG. 2, a thermally-grownoxide layer 30 and a bond coat 20 are formed on the substrate 10 andunderlie the thermal barrier coating 40. The bond coat 20 is formed onthe nickel- or cobalt-based superalloy substrate, and can thereforereact with available oxygen to form the thermally-grown oxide layer 30.Exemplary bond coat materials include an oxidation resistant alloy suchas MCrAlY, wherein M is cobalt and/or nickel, or an oxidation resistantintermetallic, such as diffusion aluminide, platinum aluminide, anactive element-modified aluminide, and combinations of the same. Anexemplary bond coat 20 ranges in thickness between about 1 and about 6mils (between about 25 and about 150 μm). The thermally-grown oxidelayer 30 is grown from the aluminum in the above-mentioned bond coatmaterials, and is ideally less than 2 μm thick.

Having described the general structure of a thermal barrier coating andexemplary underlying coating layers, FIG. 3 outlines a coatingproduction method. As mentioned previously, in some cases it may beadvantageous to include a bond coat over the substrate 10 and underlyingthe thermal barrier coating 40. If a bond coat is to be included in thecoating structure, step 50 comprises coating the substrate with asuitable metal material to form the bond coat 20. Then, step 52comprises growing an oxide layer 30 over the exposed bond coat region.Growing the oxide layer 30 can be performed by heating the substrate 10and the bond coat 20 in the presence of oxygen. The oxide layer 30 maynucleate and grow during thermal barrier coating deposition when oxygenis present. For example, an aluminum oxide layer 30 will form when thearticles are maintained between about 950° C. and about 1150° C. duringthe coating deposition, with the oxygen pressure ranging between about0.0001 and about 0.02 Torr.

Step 54 comprises depositing the ceramic thermal barrier coating innerlayer 42. A physical vapor deposition process may be used to form acolumnar, porous ceramic structure. An exemplary physical vapordeposition process for depositing the outer and inner layers 44, 42 iselectron beam-physical vapor deposition. Furthermore, if the articleincluding the substrate 10 with the bond coat 20 formed thereon istransferred into a physical vapor deposition chamber, the thermal growthoxide 30 and the entire thermal barrier coating 40 can be formed insidethe chamber without moving the article out of or into the chamber again.

The thermal barrier coating inner layer 42 is grown by mounting thearticle onto a rotating stage in a deposition chamber that includes adeposition source. The rotating stage supports the substrate above thecoating vapor source. As the article rotates, ceramic material isdeposited onto the substrate surface. More particularly, the surfacearea that is in the line of sight of the coating vapor source receivesthe ceramic material. Deposition on a surface area is interrupted whenthat area is rotated out of the line-of-sight. Tips of individualcolumnar grains of a stabilized zirconia or hafnia coating 42 aretypically slanted or pointed due to growth on crystallographic planes.Deposition shadows exist between adjacent grain tips when theline-of-sight vapor is deposited at very low angles with respect to thecomponent surface; for example, deposition shadows between grain tipsare formed when coating begins and stops during each rotation of thesubstrate over the electron beam-physical vapor source. Shadoweddeposition and associated porosity result in submicron thicknessintercolumnar gaps between the growing columnar grains. Consequently,rotation of the substrate over the vapor source cause the inner layer 42to have a columnar-grained microstructure with intercolumnar gapporosity, which provides a mechanism for strain accommodation in thickcoatings.

An exemplary method for growing the thermal barrier coating inner layer42 includes bleeding an oxygen stream into the vacuum coating chamberwhile rotating the article and performing the deposition process. Thepresence of oxygen in the chamber during deposition of the thermalbarrier coating enhances the oxygen stoichiometry of the inner layer 42,minimizes metallic atom mobility during deposition, and increases thewidth of the submicron-thickness intercolumnar gaps between the columnargrains.

Step 56 comprises depositing the thermal barrier coating outer layer 44,which is a higher density layer with respect to the thermal barriercoating inner layer 42. An exemplary method for forming the high densityouter layer 44 includes halting rotation of the article so that surfacesrequiring the dense thermal barrier coating surface layer 42 haveuninterrupted line-of-sight to the vapor source during deposition.Halting substrate rotation effectively removes the periodic depositionshadow that promotes formation of intercolumnar porosity. Further,halting substrate rotation increases radiant heating of the surfacesexposed to the vapor source, which increases oxide molecule and metallicatom mobility, which in turn promotes closure of the intercolumnar gapsbetween grains.

Density is also increased by setting the oxygen bleed pressure to lessthan 0.0001 torr while halting rotation of the article. Since ceramicssuch as zirconia and hafnia partially decompose during electron beamevaporation, the deposited coating will be oxygen deficient when theoxygen bleed is turned off. Metallic zirconium atoms diffuse faster thanzirconium oxide molecules and promote closure of intercolumnar gaps.Secondary oxides such as tantala and niobia, which are less stable invacuum than zirconia, further enhance surface diffusion and associatedclosure of intercolumnar gaps. Alumina, which has a metastable crystalstructure during deposition, may also enhance surface diffusion until itcrystallizes into the stable alpha phase or reacts with a stabilizingoxide, such as yttria, to form particles of yttrium aluminum garnet. Theresulting dense outer layer 44 protects the underlying porous thermalbarrier coating layer 42 from penetration by contaminants such as CMAS.

The oxygen bleed can be set to zero with or without rotation of thearticle on which the outer layer 44 is being deposited. When oxygendeficient deposition results in sufficient molecular and atomic mobilityto densify the outer layer 44, halting substrate rotation is optional.

In one exemplary embodiment, the outer layer 44 includes the samechemicals as the inner layer, but has a denser overall structure. In oneexemplary embodiment, the concentration of the oxide dopants in theouter layer 44 ranges between about 0.5% and about 20%, and the oxidedopant concentration in the outer layer 44 is the same as in the innerlayer 42. In another exemplary embodiment, the denser outer layer 44includes an increased concentration of a dopant such as tantala, niobiaor alumina, with respect to the inner layer 42, to enhance surfacediffusion and densification.

Before removing the article from the deposition apparatus, step 58comprises resuming the oxygen bleed for a predetermined time period torestore the oxygen stoichiometry of the outer layer 44. Oxidation of theoxygen deficient surface layer 44 causes a small volumetric expansion inan outermost region of the outer layer 44, further closing and sealingany intercolumnar gaps in the outer layer 44. This final oxidation step58 may also be performed in a post-coating heat treatment or duringengine service.

While the invention has been described with reference to a preferredembodiment, it will be understood by those skilled in the art thatvarious changes may be made and equivalents may be substituted forelements thereof without departing from the scope of the invention. Inaddition, many modifications may be made to adapt to a particularsituation or material to the teachings of the invention withoutdeparting from the essential scope thereof. Therefore, it is intendedthat the invention not be limited to the particular embodiment disclosedas the best mode contemplated for carrying out this invention, but thatthe invention will include all embodiments falling within the scope ofthe appended claims.

1. A turbine engine component, comprising: a substrate; and a thermalbarrier coating covering at least a portion of the substrate, thethermal barrier coating comprising: an inner layer having acolumnar-grained microstructure with inter-columnar gap porosity, andcomprising a stabilized ceramic material, and a substantially non-porousouter layer, covering the inner layer and comprising the stabilizedceramic material that has an oxygen deficient stoichiometry with respectto the inner layer.
 2. The turbine engine component of claim 1, whereinthe thermal barrier coating outer layer is less than about 1 mil inthickness.
 3. The turbine engine component of claim 2, wherein thethermal barrier coating outer layer is no greater than about 0.2 mil inthickness.
 4. The turbine engine component of claim 1, wherein thestabilized ceramic material in both the thermal barrier coating innerand outer layers comprise a material selected from the group consistingof oxide-stabilized tetragonal and cubic zirconia, oxide-stabilizedtetragonal and cubic hafnia, and mixtures thereof.
 5. The turbine enginecomponent of claim 1, wherein the thermal barrier coating inner layercomprises a dopant oxide at a concentration lower than a concentrationof the dopant oxide in the thermal barrier coating outer layer.
 6. Theturbine engine component of claim 5, wherein the dopant oxide has aconcentration in the outer layer ranging between about 0.5 and about 20weight %, and is selected from the group consisting of tantala, niobia,and alumina.
 7. The turbine engine component of claim 5, wherein thedopant oxide is more readily reducible than the stabilized ceramicmaterial.